The present invention relates to a gas turbine engine impeller, more particularly such an impeller having an integral annular collar extending around the periphery of the rotor disk so as to define a platform in between the plurality of blades.
The performance of modern gas turbine engines, particularly those utilized in aeronautical applications, is increased by increasing the operational temperatures, especially those at the turbine intake, as well as increasing the rotational speeds of the turbine to increase the gas volume passing through the passages defined across the turbine blades and stators.
In order to reduce the weight of the impellers, and therefore reduce the centrifugal forces acting thereon, the blades are typically mounted on rotor disks having comparatively small diameters. The weight is further reduced by fabricating the disks and blades from composite materials, or other types of materials which are resistant to high temperatures and centrifugal forces.
For reasons of manufacturing economy, the composite blades are typically made without platforms such that they have an aerodynamic profile extending substantially from their tip to their root. Once the blades are assembled to the rotor disk, platforms must be separately attached thereto. The platforms define the inner boundary of the passage through which the turbine gasses pass. The platforms also have a diameter which generally exceeds the diameter of the rotor disk in order to avoid contact between the turbine gases and the rotor disk to minimize the turbulence of the gases passing through the turbine. The platforms also serve to reduce the heat transfer from the turbine gases to the rotor disk, thereby protecting it from excessively high temperatures.
In known types of gas turbine engine impellers, the platform comprises a plurality of separate and discrete elements which are assembled to form a platform ring, the elements being also separate from the blades and affixed to the rotor disks.
U.S. Pat. No. 2,834,573 discloses a rotor construction wherein the platform ring extends between the blades and is formed from a plurality of individual, separate ring segments. Each segment extends between a pair of adjacent turbine blades.
French Patent 1,501,492 discloses a compressor impeller wherein the blades are formed without platforms and wherein the platform is formed by a plurality of individual segments extending between adjacent blades and attached to the rotor disk.
French Patent 2,073,854 describes an impeller rotor in which segments extending between adjacent blades provide mechanical damping to minimize blade vibration, but also serve as gas flow platforms. Again, these segments are attached to the rotor disk adjacent to the blade roots.
U.S. Pat. No. 4,802,824 discloses a turbine impeller wherein the blade roots are held in place within cavities defined by the rotor disk by a plurality of wedge-shaped segments which also act as platforms.
In all of the structures noted above, the separate, individual segments which constitute the platform ring must be rigidly affixed to the rotor disk because of the high centrifugal forces acting on them during the high speed operation of the gas turbine engine. The large number of separate elements increases the time required to assemble the gas turbine impeller, thereby increasing the manufacturing costs. Similarly, a great amount of time is required when the impeller must be disassembled for repairs or routine maintenance.